Bonded and tailorable composite assembly

ABSTRACT

An all-composite assembly such as a composite laminate aircraft empennage has vertical and horizontal stabilizers with differing sets of interlaminar fracture toughnesses and differing stiffnesses to improve flight characteristics. Composite laminate skins are bonded to unitized and stiffened understructure to reduce weight and improve damage containment.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is a continuation-in-part of prior U.S. patentapplication Ser. No. 13/596,989 filed Aug. 28, 2012, the entiredisclosures of which are incorporated by reference herein.

BACKGROUND INFORMATION

1. Field

The present disclosure generally relates to composite structures, anddeals more particularly with a bonded and tailorable composite assembly,such as an aircraft empennage.

2. Background

Composite materials have been integrated into airfoil designs in recentyears in order to reduce aircraft weight. For example, airfoils such asvertical and horizontal stabilizers have been designed with an outercomposite laminate skin. However, due to design limitations, interiorstructural components of these stabilizers, such as spars, stringers andchords, are predominantly fabricated from metal such as aluminum ortitanium, all secured together with metal fasteners. These metalcomponents add undesired weight to the aircraft, and are both timeconsuming and labor intensive to fabricate and assemble. In order todecrease aircraft weight and increase fuel savings, a greater use oflightweight materials such as bonded, rather than fastened composites isrequired, however known designs have limitations that make replacementof metal components with composites challenging. Moreover, theselimitations make it difficult to obtain certification of components andsubassemblies by certification authorities.

In addition to the problems discussed above, vertical and horizontalstabilizer designs that rely on extensive use of metal components arenot easily tailored to optimize a combination of flight characteristics,such as lift, bending and torsional stiffness, and discrete damagecontainment/arrestment. Also, the use of many separate spars and/orchords in these stabilizer designs may make it difficult to optimizeload transfer from the outer skins. Further, known stabilizers havingcontrol surfaces such as elevators and rudders often require theincorporation of flutter pumps or reactive masses that are connected tothe control surfaces in order to control aeroelastic stability commonlyreferred to as flutter. The use of these flutter control devices addsundesired weight and complexity to the aircraft.

Accordingly, there is a need for a composite assembly such as anaircraft empennage having unitized and bonded, all composite verticaland horizontal stabilizers that substantially reduce or eliminate theneed for metal components, including fasteners, and eliminate the needfor flutter control devices. There is also a need for a more efficientempennage with vertical and horizontal stabilizers having compositeouter skins with individually tailored interlaminar fracture toughnessesas well as stiffnesses, and an integrated all-composite gridunderstructure structure that permits tailoring the stabilizers tooptimize flight characteristics, service life durability and failsafereliability.

SUMMARY

The disclosed embodiments provide a multi-functional, tailored, andintegrated composite bonded orthotropic composite assembly, such asaircraft empennage, with a Z-stiffened grid-like compositeunderstructure. The all-composite bonded empennage has twomulti-functional aircraft horizontal stabilizers and an aircraftvertical stabilizer, each of which has uniquely designed structuralproperties with different, tailorable, interlaminar fracture toughnessmodes, improved structural elastic constants, and stiffnesses providedby the disclosed skins and bonded understructures. These qualitiesprovide the empennage with improved structural fail-safe, higherdurability and damage tolerance, higher and improved aerodynamicallybalanced lift, and substantially improved critical aerodynamic stabilityand control, with substantial reductions in aircraft weight. Anempennage incorporating a combination of a multi-performance orthotropicbonded composite laminate skin, with a Z-stiffened integrated gridunderstructure having tailorable different interlaminar fracturetoughnesses, may provide substantial improvements in fuel savings,damage containment capability, stability control, and fail-safe designof aircraft incorporating such empennages. In addition, the structuralproperties of the empennage may eliminate the need for flutter controldevices, thereby reducing aircraft weight and complexity. While anaircraft empennage is disclosed, the all bonded, composite assembly maybe used in a variety of applications, including but not limited toaerospace vehicles, marine vehicles, land vehicles, and wind drivenmachines, to name only a few. The disclosed composite assembly may alsobe used in non-vehicle related applications, such as in the buildingconstruction and other industries.

The skins each have interlaminar fracture toughness mixed Modes thatprovide increased fiber stiffness, and improved reaction to globalbending and torsion loads on control surfaces. The bonded compositeempennage structure is divided into multiple interlaminar fracture ModesI, II, III, resulting in higher stability control for upper and lowerskin surfaces that are integrally bonded to the Z-stiffenedunderstructure. “Mixed Modes” refers to the presence of a combination ofModes I, II and III that result in complex skin loading interactions.The integrated understructure reduces unnecessary high hinge loads oncritical control surfaces, and increases lift with minimum aircraftrotation while damping flutter loads on the empennage. The compositionof the bonded skins of each of the stabilizers of the empennage istailorable, having differing interlaminar fracture toughnesses in ModesI, II and III which increase the capability of the structure to containany accidental discrete damage caused by an engine explosion or impactswith foreign objects, thereby improving damage tolerance required byaircraft certifying authorities.

The understructure of the horizontal and vertical stabilizers comprisesthree main composite high modulus spars bonded to and stiffened byZ-stiffeners which have high fiber stiffness, improved differentinterlaminar elastic constants and composites having selected Mode I, IIand III properties. The Mode I, II and III properties of the verticaland horizontal stabilizers skins are separately tailored in order toprovide the empennage with higher transverse tension, shear andtorsional stiffness. These features result in a tailorable empennagehaving higher aerodynamic balanced lift and control stability duringflight, and a reduction in the amount of effort required to move controlsurfaces by up to 50%.

The composite stabilizer skins are designed with different sets ofinterlaminar fracture toughnesses in Modes I, II and III, with highfiber elastic constants, and high stiffness to increase aerodynamicbalanced higher lift, reduce maneuver loads, stabilize the attack angleto a minimum, reduce global shear and torsion loads, reduce flutter,engine thrust, and especially large out of plane hinge loads. The skinsof the vertical stabilizer have interlaminar fracture toughnesses inModes I, II and III that are different from those of the horizontalstabilizers. The disclosed bonded and integrated empennage providesself-containment of discrete damage, and extensive reduction of theinterlaminar singular edge peel loads and shear stress at eachintersection of the understructure grid and Z-stiffeners to sustainimproved aerodynamic lift, diminish aircraft rotation, diminish hingeloads, regardless of any hard-points needed to attachment lugs for thehorizontal tails, dorsal bath-tub fitting, and stiffness mismatches.

The high out-of-plane peel and lateral empennage adhesive interfacestresses are greatly minimized to balance out the aerodynamic loadsduring high turbulence and high pitch take off loads. The interfacialbonded stresses are redistributed uniformly throughout the empennageskins and understructure, thereby reducing high torsion, global & localbending loads on control surfaces hinge lugs due to turbulence duringflight, rendering the control surfaces substantially fail-safe. A highinterlaminar mode III torsion capability also increases the highvertical deflection on the horizontal stabilizers. Thus, the disclosedempennage efficiently reacts to the gust, and maneuver loads in flight,leading to a higher balanced aerodynamic lift of the aircraft & reducedvertical flutter loads.

The understructure has a moderate to high stiffness in order to reactmost wing & fuselage flight loads such as heavy bending moments,torsion, or skin in-plane shear stresses. The bonded understructure alsominimizes up and down bending due to lateral gusts or maneuver loads byredistributing the resultant loads through the bonded fail-safe joints.The stress singularities that normally develop at the bonded jointrun-outs are also drastically reduced with an adhesive taper spew at theedges of the Z-stiffened understructure cap with empennage skins.

The disclosed empennage reduces loads on the horizontal tail/elevatorhinge, side body lug joints and bathtub fittings attachments to rudderhinge joints by load redistribution across a larger bonded area, thusreducing the overall effects of concentrated fuselage side body bendingloads in the joint during flight. The empennage design also reduces theoverall aircraft weight, eliminates fasteners, and also increases theEME capability of the empennage, and thus increases pitch angle fasterand more efficiently. The all bonded composite empennage gives theaircraft an angle to the airflow which produces a higher lift on thehorizontal tails and vertical tail.

According to one disclosed embodiment, a composite assembly comprises afirst composite structure and at least one second composite structure.The first composite structure includes a composite first understructureand a composite first laminate skin bonded to the first understructure.The composite first laminate skin has a first set of pre-selectedinterlaminar fracture toughnesses. The second composite structureincludes a composite second understructure and a composite laminatesecond skin bonded to the second understructure. The composite laminatesecond skin has a second set of pre-selected interlaminar fracturetoughnesses. Each of the composite laminate first and second skins issubject to Mode I, II and III loading. The first and second sets ofinterlaminar fracture toughness are different from each other in theMode I, II and III. The torsional stiffness of the first compositestructure is greater than the torsional stiffness of the secondcomposite structure. “Torsional stiffness”, sometimes known in the artas torsional rigidity, is a measure of the ability of an elongatestructural member such as the vertical stabilizer, to resist deformationin response to an applied torque.

The first composite structure the composite assembly has a torsionalstiffness within the range of approximately 45.0 to 52.0 million poundsper square inch, and the second composite structure has a torsionalstiffness within the range of approximately 40.0 to 50.2 million poundsper square inch. The first composite laminate skin includes a Mode Iinterlaminar fracture toughness within the range of approximately 4.0 to6.5 inch-pounds per square inch, a Mode II interlaminar fracturetoughness within the range of approximately 12.0 to 15.5 inch-pounds persquare inch, and a Mode III interlaminar fracture toughness within therange of approximately 16.0 to 18.5 inch-pounds per square inch. Thesecond composite laminate skin includes a Mode I interlaminar fracturetoughness within the range of approximately 2.5 to 3.5 inch-pounds persquare inch, a Mode II interlaminar fracture toughness within the rangeof approximately 7.5 to 9.5 inch-pounds per square inch, and a Mode IIIinterlaminar fracture toughness within the range of approximately 18.0to 20.5 inch-pounds per square inch.

The first understructure comprises a plurality of longitudinallyextending composite spars, and a plurality of Z-shaped compositestiffeners extending between and bonded to the spars. Each of thecomposite spars and the Z-shaped composite stiffeners is generallyI-shaped in cross-section and includes a pair of caps. The caps of thecomposite spars and the caps of the Z-shaped composite stiffeners arebonded together. The composite assembly further comprises a plurality ofsubstantially straight, composite cross-beams respectively passingthrough the Z-shaped composite stiffeners and extending substantiallynormal to the composite spars. The composite assembly also includes aplurality of longitudinally extending, composite stringers. The elongatecomposite stringers pass through the Z-shaped composite stiffeners andare bonded to at least one of the first and second skins. The first andsecond composite structures may be stabilizers arranged to form anaircraft empennage.

According to another disclosed embodiment, a composite structurecomprises a composite laminate skin bonded to a compositeunderstructure. The understructure includes first and second,longitudinally extending composite spars and a first plurality ofZ-shaped composite stiffeners extending between and bonded to the firstand second composite spars. Each of the first and second composite sparsand the Z-shaped composite stiffeners is substantially I-shaped incross-section. The first composite spar includes a first cap, the secondcomposite spar includes a second cap, and each of the Z-shaped compositestiffeners includes a third cap bonded to each of the first and secondcaps. The understructure further includes a plurality of substantiallystraight, composite cross-beams respectively passing through theZ-shaped composite stiffeners, each of the composite cross-beamsextending substantially normal to the first and second composite spars.Each of the Z-shaped composite stiffeners includes a web having aheight, and each of the cross-beams includes opposite ends bonded to acorresponding Z-shaped composite stiffener along substantially theentire height of the web. The composite structure may further comprise athird longitudinally extending composite spar and a second plurality ofZ-shaped composite stiffeners extending between and bonded to the secondthird composite spars. The first composite spar is a front spar, thesecond composite spar is a mid-spar, and the third composite spar is anaft-spar. Each of the first and second composite spars includes firstand second spar caps, and each of the Z-shaped stiffeners includes firstand second stiffener caps. The composite laminate skin is bonded to thefirst and second spar caps and the first and second stiffener caps.

According to still another embodiment, an aircraft empennage comprises avertical stabilizer and at least one horizontal stabilizer. The verticalstabilizer has a composite laminate first skin and a composite firstunderstructure bonded to the first skin. The composite laminate firstskin is subject to Mode I, II and III loading and has a first set ofinterlaminar fracture toughnesses in each of these three modes. Thecomposite first understructure includes an integrated grid of compositespars, composite cross-beams and composite stiffeners bonded together.The at least one horizontal stabilizer includes a composite laminatesecond skin and a composite second understructure bonded to the firstskin. The composite laminate second skin is also subject to Mode I, IIand III loading and has a second set of interlaminar fracturetoughnesses in each of the three modes. The composite secondunderstructure includes an integrated grid of composite spars, compositecross-beams and composite stiffeners bonded together. The compositelaminate first skin includes a Mode I interlaminar fracture toughnesswithin the range of approximately 4.0 to 6.5 inch-pounds per squareinch, a Mode II interlaminar fracture toughness within the range ofapproximately 12.0 to 15.5 inch-pounds per square inch, and a Mode IIIinterlaminar fracture toughness within the range of approximately 16.0to 18.5 inch-pounds per square inch. The composite laminate second skinincludes a Mode I interlaminar fracture toughness within the range ofapproximately 2.5 to 3.5 inch-pounds per square inch, a Mode IIinterlaminar fracture toughness within the range of approximately 7.5 to9.5 inch-pounds per square inch, and a Mode III interlaminar fracturetoughness within the range of approximately 18.0 to 20.5 inch-pounds persquare inch. The vertical stabilizer has a torsional stiffness withinthe range of approximately 45.0 to 52.0 million pounds per square inch,and each of the horizontal stabilizers has a bending stiffness withinthe range of approximately 30.0 to 36.5 million pounds per square inch.“Bending stiffness”, sometimes referred to in the art as flexuralrigidity, is a measure of the ability of an elongate structural membersuch as each of the horizontal stabilizers 32, to resist deformation inresponse to an applied bending moment along the longitudinal axis of thestructural member. Each of the spars has a bending stiffness ofapproximately 45 million pounds per square inch. Each of the compositestiffeners is Z-shaped, and each of the composite cross-beams passesthrough one of the Z-shaped composite stiffeners. Each of the compositefirst and second understructures further includes stringers passingthrough the composite stiffeners and bonded to a corresponding one ofthe first and second skins.

According to a further embodiment, a method is provided of making acomposite structure. A plurality of composite spars and a plurality ofcomposite stiffeners are fabricated. A composite understructure isformed by bonding the composite spars and the composite stiffenerstogether. First and second composite skins are bonded to opposite sidesof the composite understructure.

The features, functions, and advantages can be achieved independently invarious embodiments of the present disclosure or may be combined in yetother embodiments in which further details can be seen with reference tothe following description and drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The novel features believed characteristic of the illustrativeembodiments are set forth in the appended claims. The illustrativeembodiments, however, as well as a preferred mode of use, furtherobjectives and advantages thereof, will best be understood by referenceto the following detailed description of illustrative embodiments of thepresent disclosure when read in conjunction with the accompanyingdrawings, wherein:

FIG. 1 is an illustration of a perspective view of an aircraft showingvarious loads affecting flight characteristics of the aircraft.

FIG. 2 is an illustration of a side view of the vertical stabilizer ofthe aircraft shown in FIG. 1, a skin being broken away in section toreveal an understructure.

FIG. 2A is an illustration similar to FIG. 2, but showing one of thehorizontal stabilizers.

FIG. 3 is an illustration of a perspective view showing additionaldetails of the Z-stiffening between two of the spars of the horizontalstabilizer shown in FIG. 2A.

FIG. 4 is an illustration of a sectional view taken along the line 4-4in FIG. 2A.

FIG. 5 is an illustration of a flow diagram of a method of making an allcomposite, bonded aircraft stabilizer.

FIG. 6 is an illustration of an overall block diagram of a structureemploying the disclosed composite assembly.

FIG. 7 is a perspective illustration of a portion of the right-side wingshown in FIG. 1, a portion of the upper wing skin removed to betterreveal the inner wing-grid-structure.

FIG. 8 is a perspective illustration of a portion of the wing shown inFIG. 1, a part of the upper wing skin broken away to reveal details ofone embodiment of the inner wing-grid structure.

FIG. 9 is a sectional view taken along the line 9-9 in FIG. 8.

FIG. 10 is an illustration similar to FIG. 8 but showing an alternateform of the wing-grid structure.

FIG. 11 is a sectional view taken along the line 11-11 in FIG. 10.

FIG. 12 is a perspective illustration of another form of the wing-gridstructure, the wing skins not shown for clarity.

FIG. 13 is a perspective illustration of a portion of a wing employinganother form of the wing-grid structure, the upper wing skin not shownfor clarity.

FIG. 14 is a diagrammatic illustration showing how standard compositecomponents may be combined to form the wing-grid structure shown in FIG.13.

FIG. 15 is a perspective illustration similar to FIG. 13, but showing afurther form of the wing-grid structure.

FIG. 16 is a perspective illustration of another form of the wing-gridstructure employing cross beam straps, the wing skins not shown forclarity.

FIG. 17 is a block flow diagram of a method of making an aircraft wing.

DETAILED DESCRIPTION

The disclosed embodiments provide an all composite bonded andintegrated, orthotropic composite assembly such as an empennage thatsubstantially eliminates the need for metallic components, includingfasteners. While an aircraft empennage embodiment will be described indetail below for illustrative purposes, it is to be understood thatother embodiments of the disclosed composite assembly are possible whichmay be used in a wide range of other applications. The empennageincludes vertical and horizontal stabilizers with composite laminateskins that are bonded to all-composite understructures. The skins of thevertical and horizontal stabilizers have properties that are tailoredrelative to each other, and which are selected to provide the empennagewith improved structural fail-safe design, greater durability, higherand improved aerodynamically balanced lift, reduced flutter loads andimproved aerodynamic stability and control, with reductions in aircraftweight. The skins of the vertical and horizontal stabilizers possessgreater toughnesses which increase the capability of the empennage tocontain any accidental, discrete damage, thus enabling it to meet damagetolerance airworthy requirements per FAA FA R 25-571e, EASAcertification, and FA R 26. Structural properties of the disclosedempennage substantially reduce flutter, and may eliminate the need forflutter control devices on control surfaces.

Referring first to FIG. 1, an aircraft 20 includes a pair of wings 22and a composite assembly 24 which may comprise a stabilizer assembly,also sometimes referred to herein as an empennage 24, attached to afuselage 26. Engines 28 provide the aircraft 20 with thrust. Theempennage 24 includes a first composite structure 30 in the form of avertical stabilizer 30, and a second composite structure 32 in the formof a pair of horizontal stabilizers 32. In some embodiments, theempennage 24 may have more than one vertical stabilizer 30. According tothe disclosed embodiments, the vertical stabilizer 30 and each of thehorizontal stabilizers 32 may be all-composite bonded structures. Thevertical stabilizer 30 includes a movable rudder 34 for controlling yawof the aircraft 20, and the horizontal stabilizers 32 include movableelevators 36 for controlling the pitch of the aircraft 20. The verticalstabilizer 30 and the horizontal stabilizers 32 are each covered withlater discussed outer composite laminate skins having pre-selectedinterlaminar fracture toughnesses in Modes, I, II and III. Interlaminarfracture toughness is measured in units of inch-pounds of force persquare inch.

As will be discussed below in more detail, the all-composite verticalstabilizer 30 is bonded to an advanced composite structural material andhas a fiber stiffness on the order of 30 MSI (million pounds per squarefoot) and pre-selected Mode I, II and III interlaminar fracturetoughnesses. Each of the all-composite horizontal stabilizers 32 is alsobonded to a composite material and has a fiber stiffness on the order of40 MSI (million pounds per square foot), and a second set ofpre-selected Mode I, II and III interlaminar fracture toughnesses. Thefirst and second sets of interlaminar fracture toughnesses respectivelyof the vertical and horizontal stabilizers 30, 32 are different fromeach other. The composite structural material from which the verticalstabilizer 30 and the horizontal stabilizers 32 are formed may compriseany of a variety of fiber reinforced resin materials such as, withoutlimitation a CFRP (carbon fiber reinforced plastic) laminate. Theconstruction and materials used in the all-composite, bonded empennage24 provide the aircraft 20 with improved balance and lift.

A variety of loads and forces act on the aircraft 20 that must bereacted in order to maintain stable and controllable flight. Forexample, as depicted in FIG. 1, forces acting on the wings 22 and thehorizontal stabilizers 32 create moments 52, 54 about the Y axis whilethe aircraft 20 experiences thrust 42 and drag 46. The lift 38 providedby the wings 22 is reacted by the weight 40 of the aircraft 20 andbalancing loads 44. The empennage 24 must also react to elevator forces48 and rudder forces 50.

In accordance with the disclosed embodiments, the vertical stabilizer 30and the horizontal stabilizers 32 each has uniquely designed structuralproperties with differing sets of pre-selected interlaminar fracturetoughnesses in Modes I, II and III, improved structural elasticconstants, and improved stiffnesses. These qualities provide theempennage 24 with improved structural failsafe, higher durability anddamage tolerance, higher and improved aerodynamically balanced lift andgreatly improved critical aerodynamic stability and control whilereducing the weight of the aircraft 20. The disclosed empennage 24reduces loads on hinge points between the horizontal stabilizers 32 andthe elevators 36, side body lug joints and bathtub fitting attachmentsto rudder hinge joints by redistributing the load across a larger bondedarea. Because the load is distributed over a larger bonded area, theoverall effects of concentrated fuselage side body bending loads on thehinge joints is reduced during flight. The design of the empennagediscussed below in more detail may provide the aircraft 20 with an angleto the airflow that produces a higher lift on vertical stabilizer 30 andthe horizontal stabilizers 32.

Attention is now directed to FIGS. 2, 2A and 3 which illustrate detailsof the structure of the vertical stabilizer 30 (FIG. 2) and thehorizontal stabilizer 32 (FIG. 2A). The vertical stabilizer 30 includesa first multi-performance, orthotropic, composite laminate outer skin68. The first outer skin 68 includes first and second skins 68 a, 68 bon opposite sides of the stabilizer 30 which cover and are bonded to afirst unitized, all composite, integrated grid understructure 66. Thefirst understructure 66 transfers loads from the skin 68 to the fuselage26 (FIG. 1). The skin 68 may comprise, for example and withoutlimitation, a composite laminate skin having multiple plies of carbonfiber epoxy.

Each of the horizontal stabilizers 32 includes a secondmulti-performance, orthotropic, composite laminate outer skin 69. Thesecond outer skin 69 includes first and second skins 69 a, 69 b onopposite sides of the horizontal stabilizer 32 which cover and arebonded to a second unitized, all composite, integrated grid,all-composite understructure 71. The second understructure 71 transfersloads from the skin 69 to the fuselage 26 (FIG. 1). The skin 69 maycomprise, for example and without limitation, multiple plies of carbonfiber epoxy.

As will be discussed below in more detail, although the first and secondunderstructures 66, 71 are similar in structural arrangement andcomponent parts, the first and second outer skins 68, 69 possessdiffering material properties which result in the vertical stabilizer30, and the horizontal stabilizer 32 having differing, but specificallypre-selected sets of interlaminar fracture toughnesses in Modes I, IIand III which tailor the empennage 24 to provide maximum performance fora given aircraft application. This tailoring of the empennage 24 mayeliminate the need for flutter control devices for control surfaces suchas the rudder 34 and the elevators 36. The selection of materialproperties for the first and second skins 68, 69 to achieve desired butdifferent interlaminar fracture toughnesses in Modes I, II and III willsometimes be referred to herein as tailoring or “tailored” stiffnesses.

More particularly, the first and second skins 68, 69 have selected,combined interlaminar fracture toughness Modes I, II and III to provideincreased fiber stiffness, and improved reaction to global bending andtorsional loads on control surfaces such as the rudder 34 (FIG. 1) andthe elevators 36. Structural composite stiffness properties are partlyderived from high elastic modulus fibers of advanced compositestructural materials used in the first skin 68 on the verticalstabilizer 30, and the second skin 69 used in each of the horizontalstabilizers 32. As previously mentioned, each of the first and secondskins 68, 69 may comprise a composite laminate such as, withoutlimitation, carbon fiber epoxy. High stiffness-to-strength ratio fibersin the structural resin of the laminate reinforce high interlaminartoughness of the structural resin with specified structural propertiesin Modes I, II & III. The Mode I property provides for load carryingcapability of stabilizers 30, 32, while the Mode II property providesfor in-plane loads and for resisting damage of the stabilizers 30, 32.The mode III property provides for twisting/torsional stiffness of thestabilizers 30, 32.

The combination of first and second composite skins 68, 69 havingdifferent interlaminar fracture toughnesses, and the stiffened,integrated grid composite understructures 66, 71 may result in areduction in the weight of the empennage 24 while improving damagecontainment capability, stability control and failsafe design. Tailoringthe interlaminar fracture toughnesses of the first and second skins 68,69 skins provides the empennage 24 with a greater ability to contain anyaccidental, discrete damage caused by for example and withoutlimitation, an engine explosion or an impact with foreign object,thereby proving damage tolerance.

The interlaminar fracture toughnesses of the first skin 68 in Modes I,II and III for the vertical stabilizer 30 are selected to be differentfrom those of the second skin 69 on the horizontal stabilizers 32. Inone embodiment, the interlaminar fracture toughnesses of the first skin68 on the vertical stabilizer 30 are approximately within the followingranges:

-   Mode I: 4.0 to 6.5 inch-pounds per square inch,-   Mode II: 12.0 to 15.5 inch-pounds per square inch,-   Mode III: 16.0 to 18.5 inch pounds per square inch.    The tension, shear and torsional stiffness of the first skin 68 on    the vertical stabilizer 30 is approximately within the range of 45.0    to 52.0 million pounds per square inch, and the bending stiffness of    the first skin is approximately within the range of 35.0 to 38.0    million pounds per square inch. The interlaminar fracture    toughnesses of the second skin 69 on each of the horizontal    stabilizers 32 are approximately within the following ranges:-   Mode I: 2.5 to 3.5 inch pounds per square inch,-   Mode II: 7.5 to 9.5 inch pounds per square inch,-   Mode III: 18.0 to 20.5 inch pounds per square inch.    The tension, shear and torsional stiffness of the second skin 69 on    each of the horizontal stabilizers 32 is approximately within the    range of 40.0 to 50.2 million pounds per square inch, and the    bending stiffness of the second skin 69 on each of the horizontal    stabilizers 32 is approximately within the range of 30.0 to 36.5    million pounds per square inch.

Selection of higher interlaminar fracture toughness in Modes II and IIIfor the skin 68 of the vertical stabilizer 30 relative to the skin 69 onthe horizontal stabilizers 32, along with its higher tension, shear andtorsional stiffness, effectively allows the vertical stabilizer 30 todampen flutter aerodynamic loads on the aircraft 20. Further, the use ofmixed Modes in the skin 68 reduces abnormal global bending effects atthe hinge points of the rudder 34 and elevators 36.

Differing interlaminar fracture toughnesses of the skin 68 on thevertical stabilizer 30 and the skin 69 on the horizontal stabilizer 32may be achieved by varying any one or more of several skin parameters,resulting in the two skins 68, 69 have varying differing stiffnesses.For example, although the ply schedule (stack) used to fabricate theskins 68, 69 respectively on the vertical and horizontal stabilizers 30,32 may both be orthotropic, the ply schedule used for the one of theskins 68, 69 may have fewer 0° plies than the ply schedule used for theother skin 68, 69, resulting in one of the skins 68, 69 being less stiffand orthotropic than the other skin 68, 69. Alternatively, the desireddifference in interlaminar fracture toughness of the two skins 68, 69may be achieved by using a different resin, using a different fibermaterial, or using different fiber diameters.

As shown in FIGS. 2, 2A and 3, the first understructure 66 of thevertical stabilizer 30 (FIG. 2) and the second understructure 71 of thehorizontal stabilizers 32 (FIG. 2A) each comprises a grid-like, allbonded arrangement of composite structural members. The understructures66, 71 each have a moderate to high stiffness in order to react mostwing and fuselage light load conditions such as heavy bending moments,torsion or skin in-plane shear stresses. “Stiffness”, also known in theart as rigidity, refers to the ability of understructure 66, 71 toresist deformation in response an applied bending and/or torsionalloads. The bonded second understructure 71 also minimizes up and downbending of the horizontal stabilizers 32 due to lateral gusts ormaneuver loads by redistributing the resultant loads through the bondedfailsafe joints between the components of the second understructure 71.

The understructures 66, 71 of the vertical stabilizer 30, and thehorizontal stabilizers 32 are generally similar or identical inconstruction and arrangement, although the size and dimensions of theirrespective component parts may vary, depending upon the particularapplication. Each of the understructures 66, 71 comprises a front spar70 at the leading edge 56 of the stabilizer 30, 32, a mid spar 72 and anaft spar 76 at the trailing edge 58 of the stabilizer 30, 32. The spars70, 72, 74 each has a relatively high elastic modulus and extend fromthe root 62 to the tip 62 of the stabilizer 30, 32, and divide theunderstructure 66, 71 into two cells 64. In other embodiments, however,the understructure 66, 71 may comprise more than three spars 70, 72, 74and more than two cells 64. In one embodiment, each of the spars 70, 72,74 may have a stiffness of approximately 45 million pounds per squareinch. In the case of the vertical stabilizer 30 shown in FIG. 2, therudder 34 is pivotally attached to the aft spar 74 by a series of lugs78 which transfer loads from the rudder 34 to the understructure 66.Similarly, as shown in FIG. 2A, the elevators 36 of the horizontalstabilizers 32 are pivotally attached to the aft spar 74 by a series oflugs which transfer loads from the elevators 36 to the understructure71. The root 60 of the spars 70, 72, 74 in each of the vertical andhorizontal stabilizers 30, 32 is attached to bathtub fittings (notshown) or similar fittings on the fuselage 26 by attachment lugs 86which may be formed of a metal, a composite or a combination of acomposite and a metal.

The understructures 66, 71 each further include a plurality oflongitudinally spaced, Z-shaped composite stiffeners 80, hereinaftercalled Z-stiffeners 80, which extend between and are bonded to the spars70, 72, 74 at spaced apart locations along the span of each of thestabilizers 30, 32. The Z-stiffeners 80 function to stiffen the spars70, 72, 74, as well as the skins 68, 69. The skin 68 of the verticalstabilizer 30 is bonded to the spars 70, 72, 74 as well as to theZ-stiffeners 80, as shown in FIG. 2. Similarly, skin 69 of thehorizontal stabilizers 32 to is bonded the spars 70, 72, 74 as well asto the Z-stiffeners 80, as shown in FIG. 2A. Loads are transferredbetween the spars 70, 72, 74 by a plurality of composite cross-beams 82which pass through the middle of the Z-stiffeners 80. The cross-beams 82effectively divide up the load transferred between the spars 70, 72, 74.The spars 70, 72, 74 and the Z-stiffeners 80 have high fiber stiffness,differing interlaminar elastic constants and mixed Modes I, II, III toprovide improved aerodynamic balanced lift and control stability duringflight, thereby reducing the effort required to move control surfacessuch as the rudder 34 and the elevators 36.

The skins 68, 69 are each further stiffened by elongate compositestringers 84 or similar stiffeners that extend in the span-wisedirection of the respective stabilizers 30, 32 and are located betweenadjacent ones of the spars 70, 72, 74. The stringers 84 are bonded tothe skins 68, 69 using a suitable structural adhesive that may be infilm or paste form. The integrated and unitized understructures 66, 71reduce unnecessary high hinge loads on critical control surfaces such asthe rudder 34 and the elevators 36, and increase lift with minimumaircraft rotation while damping flutter loads on the empennage 24.

Referring to FIGS. 3 and 4, in one embodiment, each of the spars 70, 72,74 may comprise a composite laminate such as carbon fiber epoxy formed,for example and without limitation, by bonding or co-curing two C-shapedmembers back-to-back along with the upper and lower caps. In theillustrated example, each of the spars 70, 72, 74 is an I-beam having apair of caps 94, 96 connected by a web 98, however other cross-sectionalshapes are possible. The Z-stiffener 80 is a Z-shaped beam having a pairof caps 90, 92 connected by a web 88. The Z-stiffener 80 includes twoouter legs 102, 104 bonded, and extending substantially parallel to thespars 70, 72, 74, and a diagonally extending leg 100 which is formedintegral with the outer legs 102, 104. The Z-stiffener 80 may be acomposite laminate such as carbon fiber epoxy that may be laid up,formed and cured using known techniques. The skins 68, 69 are adhesivelybonded to the caps 90, 92 of the Z-stiffeners 80, as well as to the caps94, 96 of the spars 70, 72, 74 of the respectively associatedunderstructures 66, 71.

The cross-beam 82 passes through the middle leg 100 of the Z-stiffener80 and is bonded to the web 88 at the outer legs 102, 104 of theZ-stiffener 80. The height of the cross-beam 82 is substantially equalto the height of the webs 88, 98. The cross-beam 82 as well as the web88 of the Z-stiffener 80 may have a mouse hole-like opening 106 thereinto allow pass through of each of the stringers 84. The stringer 84 mayhave any of several known cross-sectional shapes, and in the illustratedexample, is a blade type stiffener having one side 84 a thereof bondedto the skin 68. The caps 90, 92 of the Z-stiffener 80 are respectivelybonded at 95 (see FIG. 4) to the caps 94, 96 of the spars 70, 72, 74using conventional techniques, such as by applying and curing a film orpaste adhesive between the joining surfaces of the caps 90, 92, 94, 96.The edges of the caps 90, 92 of each of the Z-stiffeners 80 may have anadhesive paper spew (not shown) in order to reduce stress singularitiesthat normally develop at the bonded joint run-outs.

The composite assembly 24, such as the empennage 24 comprising theall-composite vertical stabilizer 30, and the composite horizontalstabilizers 32, may be assembled using a series of steps broadlyindicated in FIG. 5. At 108, the composite spars 70, 72, 74 arefabricated by laying up, forming and co-curing composite laminatecomponents of each of the spars 70, 72, 74 and then bonding themtogether. Similarly, at step 110, the composite Z-stiffeners 80 arefabricated by laying up, forming and co-curing composite laminatecomponents which are then bonded together. It may be possible tointegrate cross-beams 82 into the Z-stiffeners 80 as part of thefabrication of the Z-stiffeners 80. At step 112 the understructure 66 isformed by placing the Z-stiffeners 80 between the spars 70, 72, 74 usingsuitable assembly tooling (not shown). At 114, the components of theunderstructure 66, 71 are integrated into a unitized grid by bonding theZ-stiffeners 80 the spars 70, 72, 74. With components of theunderstructure 66, 71 having been bonded together, then, at step 116,the composite skins 68, 69 are respectively bonded to opposite sides ofthe all-composite understructures 66, 71.

As previously mentioned, the aircraft empennage 24 described above isonly one illustrative embodiment of the disclosed composite assembly 24.Other embodiments are possible that are suitable for use in otherapplications. Referring now to FIG. 6, a composite assembly 120employing the principles discussed above may be used on, in or form apart of a structure 118. For example and without limitation, thestructure 118 may comprise a windmill or other apparatus, and thecomposite assembly 120 may be employed to support the windmill otherapparatus. The composite assembly 120 may be similar in construction,and have features and characteristics similar to the composite assembly24 previously described.

The composite assembly 120 comprises a first composite structure 122 andat least one second composite structure 124. The first compositestructure 122 may have a construction, features and characteristics thatare similar to those of the vertical stabilizer 30 previously described.The first composite structure 122 broadly comprises a composite firstlaminate skin 126 bonded to and covering a first compositeunderstructure 128. The composite first laminate skin 126 may be similarin construction, features and characteristics to the composite outerskin 68 previously described, and possesses a first set of preselectedinterlaminar fracture toughnesses suitable for the application. Thefirst composite understructure 128 may be similar in construction,features and characteristics to the all composite, integrated grid,understructure 66 previously described and may include all compositespars 130, Z-stiffeners 132, crossbeams 134 and stringers 136.

The second composite structure 124 may be similar in construction,features and characteristics to the horizontal stabilizers 32 previouslydescribed. The second composite structure 124 includes a compositesecond laminate skin 130 possessing a second set of preselectedinterlaminar fracture toughnesses suitable for the application. Thecomposite second laminate skin 130 may be similar in construction,features and characteristics to the second composite outer skin 69previously described. The second composite structure 124 may furtherinclude a second composite understructure 138 bonded to and covered bythe composite second laminate skin 130. The second compositeunderstructure 138 may be similar in construction, features andcharacteristics to the all composite, integrated grid understructure 68previously described, and may include all composite spars 140,Z-stiffeners 142, crossbeams 144 and stringers 146.

Principles and features of the composite assembly 24, 120 describedabove may be incorporated into or used in combination with otherembodiments and structures, such as, without limitation, the aircraftwing 22 described below and illustrated in FIGS. 1 and 7-16. Theaircraft wing 22 shown in FIGS. 1 and 7-16 utilizes wing skin-griddifferential features to improve wing-fuselage structural performance,and reduce manufacturing costs through lighter weight bonded designs.The disclosed wing 22 may also reduce part count, may reduce oreliminate corrosion and may provide a higher structural margin ofsafety. The bonded aircraft wing 22 exhibits increased wing designefficiency, is extremely light weight and provides fuel savings, whilereducing or eliminates the need for fasteners to fasten the wing skinsto the inner wing-grid and the spars. The wing 22 has the ability tocontain discrete damage, such as that caused by engine explosion.

In one exemplary embodiment, the bonded composite aircraft wing 22 mayinclude a composite inner wing-grid structure 152 (hereinafter sometimesreferred to as a wing-grid or wing-grid structure), and upper and lowercomposite wing skins 25, 27 that may be specifically tailored to satisfydifferent load cases, such as higher lift, loads during maneuvers, upand down bending, shear and torsional loads, lateral gusts, and enginethrust. The wing's fail-safe bonded inner wing-grid structure 152provides self-containment in the event of discrete damage andsubstantial reduction of the interlaminar singular peel and shear stressat intersections of wing-grid spars and grid cross beams which form thewing-grid structure 152. Adhesives may be used to bond the wing-gridspars and/or the grid cross-beams to upper and lower composite wingskins. The upper and lower wing skins 25, 27 may have differinginterlaminar fracture toughnesses that include a graduated stiffnessreacting wing loads. Unitized constant interface bonded propertiesthroughout the wing-grid mitigate torsional loads and bending due toturbulence.

FIG. 7 is a perspective illustration of one of the wings 22 shown inFIG. 1, a portion of the upper composite wing skin 25 being broken awayto reveal an inner composite wing-grid structure 152 which will bedescribed in more detail below. The composite wing-grid structure 152comprises a grid of intersecting composite structural support memberswhich may be cocured, and is bonded to the inner surfaces 148, 150respectively of the upper and lower composite wing skins 25, 27 using asuitable adhesive which may be in film or paste form. Thus, the upperand lower wing skins 25, 27 are attached to the wing-grid structure 152by bonded joints, thereby obviating the need for discrete fasteners.

The wing-grid structure 152 comprises a plurality of wing-grid spars 158extending in the span-wise or X-direction from the root 29 to the tip 31of the wing 22, and a plurality of intersecting composite supportmembers, hereinafter referred to as grid cross beams 160, extending inthe chord-wise or Y-direction, traverse to the wing-grid spars 158. Thewing-grid spars 158, grid cross beams 160 and upper and lower skins 25,27 may form a wing box that includes a leading edge grid spar 158 a, anda trailing edge grid spar 158 b to which leading and trailing edgeassemblies (not shown) are respectively attached.

In some embodiments, the leading and trailing edge grid spars 158 a, 158b may be larger and/or stiffer that the other wing-grid spars 158(sometimes referred to as mid-body spars) in order to transfer loadsbetween the wingbox and the leading and trailing edge assemblies. Thewing-grid spars 158 may be substantially uniform in cross section andother characteristics along their respective lengths.

The upper and lower composite wing skins 25, 27 respectively havediffering interlaminar fracture toughnesses in Modes I, II and IIrespectively, resulting in the upper and lower wing skins 25, 27 havingdiffering stiffnesses that are specifically tailored to meet both staticand dynamic global loads of a particular aircraft application. Theinterlaminar fracture toughnesses of the upper and lower wing skins 25,27 may be selected such that when the aircraft 20 (FIG. 1) is on theground, the upper wing skin 25 is in tension and the lower wing skin 27is in compression, but during flight, the upper and lower wing skins 25,27 are respectively in compression and tension.

Employing different interlaminar fracture toughnesses of composite wingskins 25 and 27 in combination with the composite wing-grid structure152 better distributes wing loads during flight over a wider structuralarea, and may reduce or eliminate the need for structural chord webstypically used in traditional wing structures, while minimizing thenumber of spars required. Moreover, the use of upper and lower wingskins 25, 27 having differing interlaminar fracture toughnesses incombination with the wing-grid structure 152 better reacts a variety offorces applied to the wings 22 including bending moments, torsion, shearstresses, up and down bending due to lateral gusts or maneuver loadsduring flight.

Each of the upper and lower composite wing skins 25, 27 respectively, isorthotropic and comprises a stack of laminated layers/plies offiber-reinforced resin materials, such as carbon fiber epoxy, havingrelatively high strength-to-weight ratios. Each of the plies maycomprise unidirectional reinforcing fibers of a desired angularorientation. For example, each of the wing skins 25, 27 may comprisemultiple laminated plies respectively having fiber orientations of 0°,45° and 90°. The 0° plies are generally oriented in the span-wise or Xaxis direction, while the 90° plies are oriented in the chord-wise or Yaxis direction. The 45° plies included in the ply stack react in-plane,off angle loads and function to lower the Poisson's ratio effect.

Differing interlaminar fracture toughnesses of the upper and lower wingskins 25, 27 may be achieved by varying any one or more of several wingskin parameters, resulting in the upper and lower wing skins 25, 27having differing stiffnesses. For example, although the ply schedule(stack) used to fabricate the upper and lower wing skins 25, 27 may bothbe orthotropic, the ply schedule used for the upper wing skin 25 mayhave fewer 0° plies than the ply schedule used for the lower wing skin27, resulting in the upper wing skin 25 being less stiff and orthotropicthan the lower wing skin 27. Alternatively, the desired difference ininterlaminar fracture toughness of the upper and lower skins 25, 27 maybe achieved by using a different resin, using a different fibermaterial, or using a different fiber diameter in the upper wing skin 25,compared to that used in the lower wing skin 27. Depending on theapplication, the ply schedules for the upper and lower wing skins 25, 27may or may not vary layer-by-layer in either the span-wise or chord-wisedirections. The use of wing skins 25, 27 having differing interlaminarfracture toughnesses and stiffnesses allows the bending, torsion andvertical deflection of the wings 22 to be tailored in a manner thatoptimizes distribution of the wing lift, thereby increasing overall winglift.

Upper composite wing skin 25 may have, in one typical embodiment, aninterlaminar fracture toughness of about 3.0 in-lbs/in² to about 5.0in-lbs/in² in Mode I, about 4.5 to about 7.0 in Mode II and about 7.5 toabout 8.5 in Mode III. These ranges of interlaminar fracture toughnessfacilitate the wing's ability to react to different wing flight loadsdue to independent structural composite bending, torsion, and stiffnesscapability, while retarding or arresting the propagation of cracks. Thelower composite wing skin 27 has a higher interlaminar fracturetoughnesses than the upper wing skin 25. For example, the lowercomposite wing skin 27 may have an interlaminar fracture toughness ofabout 4.5 in-lbs/in² to about 6.5 in-lbs/in² in Mode I, 5.5 in-lbs/in²to about 8.0 in-lbs/in² in Mode II, and about 8.5 in-lbs/in² to about12.0 in-lbs/in² in Mode III. These ranges of interlaminar fracturetoughnesses provide the lower wing skin 27 with unique structuraltension-shear-stiffness characteristics that better react to wingbending and torsion loads.

Upper composite wing skin 25 is formed with a lower interlaminarfracture toughness than lower composite wing skin 27 such that duringflight, upward bending of the wing places the upper composite wing skin25 in compression while the bottom wing skin 27 is in tension. Inaddition, the composite wing-grid structure 152 provides rigidity toupper and lower wing skins 25 and 27 during flight. In otherembodiments, upper composite wing skin 25 may have an interlaminarfracture toughness of greater than or lower than 4.0 in-lbs/in², andlower composite wing skin 27 may have an interlaminar fracture toughnessof greater than or lower than 6.0 in-lbs/in², where lower wing 27 has aninterlaminar fracture toughness greater than the interlaminar fracturetoughness of upper wing skin 25.

Structural composite stiffness properties are partly derived from highmodulus fibers of advanced composite structural materials for upper andlower wing skins 25 and 27. High stiffness-to-strength ratio fibers inthe structural resin reinforce high interlaminar toughness of thestructural resin with specified structural properties in modes I, II &III critical wing load cases, respectively. The mode I property providesfor load carrying capability of wings 22, while the mode II propertyprovides for in-plane loads and for resisting damage of wings 22. Themode III property provides for twisting/torsional rigidity of wings 22.

Upper composite wing skin 25 structural properties have moderately highmode I and mode II interlaminar fracture toughness. The mode Istructural property is designed to increase the load carrying capabilityof upper composite wing skin 25 under normal loads induced by bending,and compression induced during take-off and in flight. The mode IIinterlaminar toughness property for upper composite wing skin 25 isdesigned to take more in-plane shear loads due to bending and torsion,thus increasing the capability of the wing to sustain higher aerodynamicloads.

The structural composite properties of the lower composite wing skin 27are designed to have higher mode I, II and III structural interlaminartoughnesses compared to the corresponding properties of the upper wingskin 25. These structural properties are selected to increase thecapability to carry global heavy interlaminar tension and in-plane shearloads in lower composite wing skin 27 induced by up-bending. The modeIII structural interlaminar toughness of the lower composite wing skin27 is designed to increase the capability of the wing to react to thetwisting moment at the thick inboard side of wings 22 due to heavyfuselage loads. Additionally, an increase in the mode III property,which is the twisting/torsional rigidity property, results in higherlift and produces a balanced twisting angle of the wing.

The wing-grid structure 152 discussed above may be implemented using anyof a variety of composite structural configurations employing wing-gridspars 158 that are reinforced and/or are stabilized by grid cross beams160. For example, referring to FIGS. 8 and 9, the wing-grid spars 158may each comprise a composite I-beam 159 having an upper cap 164 and alower cap 166 joined together by a web 162. The upper caps 164 areadhesively bonded to the inner surface 148 of upper wing skin 25 along abond line 165, and the lower caps 166 are adhesively bonded to the innersurface 150 of the lower wing skin 27 along a bond line 167. In thisexample, the grid cross beams 160 comprise individual, substantiallyflat composite webs 168 that extend between adjacent ones of the I-beams159 and are adhesively bonded along bond lines 163 to be webs 162 of thewing-grid spars 158. In alternative embodiments, the I-beams 159 and thegrid cross beam webs 168 may be laid up and co-cured.

Attention is now directed to FIGS. 10 and 11 which illustrate analternate form of the wing-grid structure 152. In this example, thecomposite I-beams 159 forming the wing-grid spars 158 are reinforced andstabilized by grid cross beams 160 which are also I-shaped in crosssection. Each of the grid cross beams 160 comprises upper and lower caps170, 172 joined together by a web 168. The upper and lower caps 170, 172respectively are substantially coplanar with the corresponding upper andlower caps 164, 166 of the wing-grid spars 158, thereby providing alarger bond area between the wing-grid structure 152 and the upper andlower wing skins 25, 27, compared to the embodiment shown in FIGS. 8 and9. In some embodiments, the grid cross beams 160 may be laid up andco-cured with the wing-grid spars 158, while in other embodiments theymay be adhesively joined together in a secondary bonding operation.

The upper caps 164, 170 are adhesively bonded to the inner surface 148of the upper wing skin 25 along a bond line 169 in a secondary bondingoperation. Similarly, the lower caps 166, 172 are adhesively bonded tothe inner surface 150 of the lower wing skin 27 along a bond line 171,also in a secondary bonding operation. In the example shown in FIGS. 10and 11, the grid cross beams 160 are substantially aligned in thechord-wise direction of the wing 22, however in other embodiments thegrid cross beams 160 may be staggered or offset from each other in thespan-wise direction of the wing 22.

FIG. 12 illustrates still another embodiment of the wing-grid structure152, the upper and lower composite wing skins 25, 27 not shown forclarity. The grid cross beams 160 comprise upper and lower,substantially flat, composite cross beam straps 174, 176 which may belongitudinally spaced apart in the span-wise direction of the wing 22,and which extend transversely across the wing-grid spars 158. In theillustrated example, the wing-grid spars are only generically indicatedby the numeral 158. The cross beam straps 174, 176 may be laid up overand cocured with the wing-grid spars 158. The cross beam straps 174, 176have outer, flat bonding surfaces 177 to which the upper and lower wingskins 25, 27 may be adhesively bonded along bond lines 171 in asecondary bonding operation. Alternatively, the cross beam straps 174,176 may be laid up over and cocured with the wing-grid spars 158.

FIG. 13 illustrates yet another embodiment of the wing-grid structure152 which employs I-beam type wing-grid spars 158, similar to theI-beams 159 previously discussed in connection with FIGS. 8-11. In thisexample, the I-beam shaped wing-grid spars 158 are further stiffened andstabilized by composite laminate side webs 178. In order to stiffen thewing-grid structure 152 in the chord-wise direction, each of thewing-grid spars 158 is provided with a center stiffener 179 extending inthe X-Y plane. The center stiffeners 179 extend through the webs 162 andinto the side webs 178, and function as the grid cross beams 160previously described, to provide the wings 22 with the desired torsionalrigidity and shear strength. Similar to the embodiments shown in FIGS.8-11, the upper and lower caps 164, 166 are bonded directly to the upperand lower wing skins 25, 27 respectively, resulting in transmission ofloads through bond lines (not shown) between the wing skins 25, 27 andthe wing-grid structure 154.

FIG. 14 illustrates one assembly that may be employed to fabricate thewing-grid structure 152 shown in FIG. 13, using simple pre-formedcomposite laminate segments. Four preformed C-shaped segments 182 arepre-positioned against a center web segment 190. Side web segments 180,which form side webs 178 (FIG. 13), are then positioned over the outersides of the C-shaped segments 182. Cap segments 186 are placed over theassembled C-shaped segments 182, thereby forming a layup assembly thatcomprises all of the elements of the wing-grid structure 152 shown inFIG. 13. The composite laminate segments shown in FIG. 14 are laid up insequence and then co-cured.

FIG. 15 illustrates another embodiment of the wing-grid structure 152,similar to that shown in FIG. 13, with the exception that the centerstiffener 179 extends continuously in the chord-wise direction betweenthe wing-grid spars 158, as well as continuously in the span-wisedirection of the wing 22. In some forms, the center stiffener 179 mayhave gaps therein between adjacent ones of the wing-grid spars in orderto reduce wing weight. For example, the center stiffener 179 maycomprise a series of individual, spaced apart center stiffener strapsthat pass through the wing-grid spars 40.

FIG. 16 illustrates another variation of the wing-grid structure 152that combines features of the embodiments previously discussed inconnection with FIGS. 12 and 13. In this embodiment, the grid crossbeams 160 are formed by upper and lower cross beam straps 174, 176 whichare bonded to or co-cured with the wing-grid spars 158. The upper andlower cross beam straps 174, 176 may be spaced apart in the span-wisedirection, similar to the embodiment shown in FIG. 12, or may becontinuous or semi-continuous in the span-wise direction of the wing 22.The upper and lower cross beam straps 174, 176 may have a thicknesssufficient to react transverse shear loads and bending in the chord-wisedirection.

FIG. 17 broadly illustrates the steps of a method 192 of making anaircraft wing 22 of the type previously described. Beginning at step194, upper and lower composite wing skins 25, 27 are formed whichrespectively have differing interlaminar fracture toughnesses andstiffnesses in Modes I, II and III, as well as different elasticconstants. As previously mentioned, differing interlaminar fracturetoughnesses and stiffnesses may be imparted to the wing skins 25, 27 byvarying the one or more material characteristics, such as the number ofplies having a particular ply orientation, such as 0° plies, in theskin. Following ply layup of each of the wing skins 25, 27, they areeach cured. At step 196, an inner wing-grid structure 152 is formed bylaying up the elements comprising the wing-grid spars 158 and the gridcross beams 160 previously discussed, using suitable layup and assemblytooling. Following layup, the wing-grid spars 158 and the grid crossbeams 160 are co-cured together to form a fully consolidated andintegrated wing-grid structure 152. At step 198, the pre-cured upper andlower composite wing skins 25, 27 are adhesively bonded to the pre-curedwing grid structure 152, in a secondary bonding operation, to form acompleted wing box. At step 202, leading edge and/or trailing edgeassemblies may be attached to the wing box, as desired.

The description of the different illustrative embodiments has beenpresented for purposes of illustration and description, and is notintended to be exhaustive or limited to the embodiments in the formdisclosed. Many modifications and variations will be apparent to thoseof ordinary skill in the art. Further, different illustrativeembodiments may provide different advantages as compared to otherillustrative embodiments. The embodiment or embodiments selected arechosen and described in order to best explain the principles of theembodiments, the practical application, and to enable others of ordinaryskill in the art to understand the disclosure for various embodimentswith various modifications as are suited to the particular usecontemplated.

What is claimed is:
 1. A composite assembly, comprising: a firstcomposite structure, the first composite structure including a compositefirst understructure and a composite first laminate skin bonded to thefirst understructure, the composite first laminate skin having a firstset of pre-selected interlaminar fracture toughnesses, wherein the firstunderstructure comprises: a plurality of longitudinally extendingcomposite spars, and a plurality of Z-shaped composite stiffenersextending between and bonded to the spars, and a plurality oflongitudinally extending, composite stringers passing through theZ-shaped composite stiffeners and bonded to at least one of thecomposite first laminate skin and composite second laminate skin; and atleast one second composite structure, the second composite structureincluding a composite second understructure and a composite laminatesecond skin bonded to the second understructure, the composite laminatesecond skin having a second set of pre-selected interlaminar fracturetoughnesses.
 2. The composite assembly of claim 1, further comprising:each of the composite first laminate skin and the composite laminatesecond skin being subject to Mode I, II and III loading, and the firstset of pre-selected interlaminar fracture toughness and the second setof preselected interlaminar fracture toughness differ from each other inthe Modes I, II, and III loading.
 3. The composite assembly of claim 1,wherein: the first composite structure comprises a first torsionalstiffness, and the second composite structure comprises a secondtorsional stiffness, the first torsional stiffness being greater thanthe second torsional stiffness.
 4. The composite assembly of claim 3,wherein: the first torsional stiffness comprises a range ofapproximately 45.0 to 52.0 million pounds per square inch, and thesecond torsional stiffness comprises a range of approximately 40.0 to50.2 million pounds per square inch.
 5. The composite assembly of claim1, wherein the first set of pre-selected interlaminar fracturetoughnesses of the composite laminate first skin of the first compositestructure comprises: a Mode I interlaminar fracture toughness within arange of approximately 4.0 to 6.5 inch-pounds per square inch, a Mode IIinterlaminar fracture toughness within a range of approximately 12.0 to15.5 inch-pounds per square inch, and a Mode III interlaminar fracturetoughness within a range of approximately 16.0 to 18.5 inch-pounds persquare inch.
 6. The composite assembly of claim 5, wherein the secondset of pre-selected interlaminar fracture toughnesses of the compositelaminate second skin of the second composite structure comprises: a ModeI interlaminar fracture toughness within a range of approximately 2.5 to3.5 inch-pounds per square inch, a Mode II interlaminar fracturetoughness within a range of approximately 7.5 to 9.5 inch-pounds persquare inch, and a Mode III interlaminar fracture toughness within arange of approximately 18.0 to 20.5 inch-pounds per square inch.
 7. Thecomposite assembly of claim 1, wherein: each of the composite spars andthe Z-shaped composite stiffeners is generally I-shaped in cross-sectionand includes a pair of caps, and the caps of the composite spars and thecaps of the Z-shaped composite stiffeners are bonded together.
 8. Thecomposite assembly of claim 1, further comprising: a plurality ofsubstantially straight, composite cross-beams respectively passingthrough the Z-shaped composite stiffeners and extending substantiallynormal to the composite spars.
 9. The composite assembly of claim 1,where the first composite structure and the second composite structureare arranged to form an aircraft empennage.
 10. The composite assemblyof claim 1, wherein: the first composite structure is an aircraftvertical stabilizer and the second composite structure is an aircrafthorizontal stabilizer.
 11. An aircraft empennage, comprising: a verticalstabilizer having a composite laminate first skin and a composite firstunderstructure bonded to the first skin, such that the composite firstunderstructure comprises a first integrated grid that comprises firstcomposite spars, first composite cross-beams, and first compositestiffeners bonded together, wherein the composite laminate first skincomprises: a Mode I interlaminar fracture toughness within a range ofapproximately 4.0 to 6.5 inch-pounds per square inch; a Mode IIinterlaminar fracture toughness within a range of approximately 12.0 to15.5 inch-pounds per square inch; and a Mode III interlaminar fracturetoughness within a range of approximately 16.0 to 18.5 inch-pounds persquare inch; and a pair of horizontal stabilizers, each of thehorizontal stabilizers including a composite laminate second skin and acomposite second understructure bonded to the first skin, the compositesecond understructure comprising a second integrated grid that comprisessecond composite spars, second composite cross-beams, and secondcomposite stiffeners bonded together.
 12. The aircraft empennage ofclaim 11, wherein the composite laminate second skin comprises: a Mode Iinterlaminar fracture toughness within a range of approximately 2.5 to3.5 inch-pounds per square inch, a Mode II interlaminar fracturetoughness within a range of approximately 7.5 to 9.5 inch-pounds persquare inch, and a Mode III interlaminar fracture toughness within arange of approximately 18.0 to 20.5 inch-pounds per square inch.
 13. Theaircraft empennage of claim 12, wherein the vertical stabilizercomprises a torsional stiffness in a range of approximately 45.0 to 52.0million pounds per square inch.
 14. The aircraft empennage of claim 13wherein each of the horizontal stabilizers comprises a bending stiffnessin a range of approximately 30.0 to 36.5 million pounds per square inch.15. The aircraft empennage of claim 11, wherein each of the sparscomprises a bending stiffness of approximately 45 million pounds persquare inch.
 16. The aircraft empennage of claim 11, wherein: each ofthe composite stiffeners comprises a Z-shape, and each of the compositecross-beams passes through one of a Z-shaped composite stiffener.
 17. Anaircraft empennage, comprising: a vertical stabilizer, the verticalstabilizer including a composite first understructure and a compositelaminate first skin bonded to the first understructure, the compositelaminate first skin being subject to Mode I, II and III loading andhaving a first set of interlaminar fracture toughnesses in Modes I, II,and III; and at least one horizontal stabilizer, the horizontalstabilizer including a composite second understructure and a compositelaminate second skin bonded to the second understructure, the compositelaminate second skin being subject to Mode I, II and III loading andhaving a second set of interlaminar fracture toughnesses in Modes I, II,and III that are lesser in value than the first set of interlaminarfracture toughnesses.
 18. The aircraft empennage of claim 17, whereinthe composite laminate first skin comprises: a Mode I interlaminarfracture toughness within a range of approximately 4.0 to 6.5inch-pounds per square inch; a Mode II interlaminar fracture toughnesswithin a range of approximately 12.0 to 15.5 inch-pounds per squareinch; and a Mode III interlaminar fracture toughness within a range ofapproximately 16.0 to 18.5 inch-pounds per square inch.
 19. The aircraftempennage of claim 17, wherein the vertical stabilizer comprises atorsional stiffness in a range of approximately 45.0 to 52.0 millionpounds per square inch.
 20. The aircraft empennage of claim 17, whereinthe horizontal stabilizer comprises a torsional stiffness in a range ofapproximately 40.0 to 50.2 million pounds per square inch.
 21. Theaircraft empennage of claim 17, wherein at least one understructurecomprises a Z-shaped composite stiffener.